Combustion chamber with internal jacket made of a ceramic composite material and process for manufacture

ABSTRACT

A combustion chamber, in particular for a rocket drive, comprises at least one jacket made of a composite material with a ceramic matrix. The composite material contains a fibrous structure made of carbon-containing fibers, and the fibrous structure comprises layers of fibers that form a three-dimensional matrix.

BACKGROUND AND SUMMARY OF THE INVENTION

This application claims the priority of German Patent Document DE 101 26926.9, filed on Jun. 1, 2001, the disclosure of which is expresslyincorporated by reference herein.

This invention relates to a combustion chamber, in particular for arocket drive, exhibiting at least one jacket made of a compositematerial with a ceramic matrix, wherein the composite material containsa fibrous structure made of carbon-containing fibers, as well as aprocess for manufacturing a combustion chamber, in particular for arocket drive, wherein at least one jacket is made of a compositematerial with a ceramic matrix, wherein the composite material is madeon the basis of a fibrous structure made of carbon-containing fibers andwherein silicon is fed to the fibrous structure in order to form asilicon carbide matrix. The invention is applicable, in principle, notonly to the special application of rocket drive engineering but also toother fields, such as aircraft engineering, and, in principle, also toother combustion chambers and furnaces.

The prior art discloses combustion chambers, in particular for rocketdrives, where composite materials with a ceramic matrix are used. Thedisclosure of DE 198 58 197 describes a drive with a combustion chamber,provided with an internal jacket and insulation made of carbonfiber-reinforced silicon carbide (C/SiC). DE 197 30 674 discloses acombustion chamber, in particular for a rocket drive, provided with aninternal jacket made of a fiber-reinforced ceramic material or graphiteand an external jacket made of a fiber-reinforced ceramic material. DE37 34 100 describes a combustion chamber for an aircraft, wherein aporous, woven internal wall made of silicon carbide fibers and a wovenexternal wall made of silicon carbide fibers are provided.

A significant problem with the prior art construction method of thecombustion chambers with at least one jacket made of a compositematerial with a ceramic matrix is that, for the ceramic matrixcomposites (CMC) materials or comparable composite materials used atthat time, two-dimensional (2D) structures, in particular 2D C/SiC or 2Dcarbon fiber reinforced carbon (C/C), are used. These structures exhibitonly negligible interlaminar shear strength (ILS) in relation to theshear forces, as can be the case especially in the thrust direction of arocket drive. For example, U.S. Pat. No. 6,197,411 describes suchtwo-dimensional structures for composite materials with a metal matrix.The stress conditions that can occur in the combustion chamber and thatare induced by corresponding thermal loads can result in delaminationbetween individual layers of the CMC structures, a factor that canimpair the operability of the combustion chamber and can result in thefailure of the combustion chamber.

For special applications of the rocket drives, mechanical stresses suchas vibration, bending strain, acoustical loads, that can occurespecially at launch time, must be taken into account. To this end, itis necessary that the structure of the drive exhibit adequate ability toabsorb any loads and at the same time demonstrate adequate elasticdeformation or ductility. The ceramic external structures, known fromthe prior art, do in fact exhibit high rigidity, but also poor expansionand deformation properties. In addition, the external structure must begas-tight. The manufacture of gas-tight fiber-reinforced compositeceramics, for example, C/SiC or C/C, can be realized only at a very highcost according to the prior art methods.

Therefore, one object of the present invention is to provide acombustion chamber and a method for manufacturing a combustion chamber,in particular for a rocket drive, which overcome the drawbacks of theprior art.

According to one embodiment, the invention comprises a combustionchamber, in particular for a rocket drive, exhibiting at least onejacket made of a composite material with a ceramic matrix, wherein thecomposite material contains a fibrous structure made ofcarbon-containing fibers. In addition to applications in rocket driveengineering, such a combustion chamber can be used for combustionchambers of aircraft technology or other combustion chambers orfurnaces. According to the invention, the fibrous structure consists oflayers of fibers, forming a three-dimensional matrix. Thus, there is nolonger a two-dimensional configuration of the individual fibers or thelayers of fibers, but rather the fibers or the layers of fibers arearranged in a defined way in the form of a three-dimensional matrix. Inthis respect, the fibers or layers of fibers are suitably interconnectedin order to guarantee a defined matrix structure. Such a structuresignificantly improves the stability of the composite material withrespect to shear stresses. In addition, the goal with such a definedthree-dimensional configuration of fibers is to improve and optimize theadjustability of the thermal conductivity of the composite material.

In particular, it can be provided that the fibrous structure beconstructed from first, second and third layers of fibers, wherein thefibers of the first layers extend in a first direction in space; thefibers of the second layers extend in a second direction in space; andthe fibers of the third layers extend in a third direction in space; andwherein the individual layers penetrate each other at least partially.Thus, the individual layers can comprise, for example, fibers or bundlesof fibers that are arranged parallel to each other, wherein the fibersor bundles of fibers of each layer are separated from each other so thatfibers or bundles of fibers of another layer, extending in anotherdirection in space, can be disposed in the resulting spaces. Thisfeature offers the possibility of mutual interpenetration of theindividual layers.

Furthermore, it can be provided that the individual layers of thefibrous structure be interconnected by means of textile technology. Forexample, the individual layers can be interwoven or sewn together. Thisfeature represents a simple and effective way to interconnect individuallayers of fibers or bundles of fibers.

In the case of previous combustion chambers, it was provided thatsilicon carbide fibers be used. However, they have a drawback becausethey remain stable only up to temperatures of approximately 1,200° C.For higher temperature ranges, for example at 1,500° C. or above, as canoccur in a rocket drive, such fibers are inappropriate or they areappropriate only with the provision of additional protective measures.Therefore, for applications at higher operating temperatures exceeding1,500° C., the fibrous structure made of carbon fibers is preferred.

A composite material that contains silicon carbide can be provided as aspecial composite material. This silicon carbide can also be formed, atleast to some degree, through a reaction of silicon with the fibrousstructure.

For applications of the combustion chamber under high loads, asparticularly in a rocket drive, or in a correspondingly low load-bearingstructure of the jacket made of composite material, it is provided thatthe jacket made of composite material be enveloped by a load-bearingexternal jacket. This external jacket serves to support the compositematerial jacket, which then forms a corresponding internal jacket. Toachieve adequate flexibility precisely for applications in a rocketdrive, said external jacket is made preferably of a metal material, suchas nickel, copper, or a nickel- and/or copper-based alloy, or a suitablesteel.

If, as a function of the chosen type of materials and structures of theinternal jacket made of composite material and of the external jacket,the result is that these two jackets have a significantly differentthermal expansion coefficient, then an intermediate layer, whose thermalexpansion coefficient is between that of the external jacket and that ofthe composite material jacket, can be provided to prevent stresses oreven cracks between the external jacket and the composite materialjacket. For example, a composite material with a metal matrix can beprovided as the intermediate layer, but any other suitable material canalso be provided. The intermediate layer of metal-matrix compositematerial can be affixed on the composite material jacket. To achieve abalance between the thermal expansion coefficients that is as simple andeffective as possible, it is preferably provided that the metal matrixcontains the same metal material as the external jacket. The externaljacket can be affixed on the intermediate layer.

The invention also comprises a process for manufacturing a combustionchamber, in particular for a rocket drive, wherein at least one jacketis made of a composite material with a ceramic matrix, wherein thecomposite material is based on a fibrous structure made ofcarbon-containing fibers, and wherein the silicon is fed to the fibrousstructure for the purpose of forming a silicon carbide matrix. Theinvention provides that the fibrous structure be produced as athree-dimensional matrix comprising layers of fibers. Thus, theaforementioned advantages of increased stability, with respect to theshear stresses, and better adjustability of the thermal conductivity,are achieved in a simple and defined manner.

Channel-shaped cooling structures can be provided in the area of thefibrous structure. These cooling structures can be produced in manyways. The layers of the fibrous structure can be arranged in such amanner that channel-shaped spaces remain on the surface of the fibrousstructure and/or in the fibrous structure. In this way, the coolingstructures are already defined by way of the fibrous structure itself, afeature that has the advantage that the fibers are not destroyed bysubsequent treatment steps. It can also be provided that channelstructures be not worked into the surface of the composite material byway of mechanical treatment operations until after the compositematerial has been fabricated. This manufacturing method is relativelysimple to carry out. However, channel-shaped contracting bodies, whichdecompose or dissolve in the course of manufacturing the compositematerial or thereafter and open the corresponding channels in thecomposite material, can also be placed, for example, in the fibrousstructure within the framework of manufacturing the fiber matrix.

An additional seal and load-bearing reinforcement of the channelstructures can be produced in that at least those surface areas of thecomposite material that exhibit the channel structures are coated with ametal material. Metal-lined channel structures are then produced thatbetter fulfill the requirements imposed on the coolant, flowing throughthe channels, at least for specific areas of application.

If channel structures are provided on the fibrous structure and otherstructures are supposed to be produced on this fibrous structure, thenit can be provided that channel-shaped contracting bodies be disposed onthe surface of the fibrous structure, precisely in the channelstructures present there. Then they can be decomposed or dissolved, asdescribed above, in the course of manufacturing the composite materialor thereafter and open the corresponding channels in the compositematerial.

For applications in rocket drives, a load-bearing external jacket can beaffixed on the jacket made of composite material. For the aforementionedreasons, it can be provided that an external jacket made of a metalmaterial be affixed. Such an external jacket can be affixed by way ofelectroplating, soldering, welding techniques, or other known methods.

There are, in principle, various alternatives for composite materialswith a metal matrix that can be used for the intermediate layer. Theirmanufacture is known to some degree. See, for example, U.S. Pat. No.6,197,411. However, an improved method for manufacturing such acomposite material, in particular for forming an intermediate layerbetween the internal jacket and the external jacket of a combustionchamber, would be an improvement in the art, and such an improvement isprovided by the present invention. First, a fibrous structure is affixedon the composite material jacket, and then a metal material is depositedon the fibrous structure with simultaneous infiltration of the fibrousstructure with the metal material. Thus, in one working step, acomposite material with metal matrix can be produced as the intermediatelayer, and an external jacket can be produced on the intermediate layer.The metal material can be deposited in any suitable form, for example,from a liquid or gaseous phase. However, the metal material can also bedeposited by way of an electroplating process. In particular nickel orcopper or a nickel- and/or copper-containing alloy can be used here asthe metal material.

The invention also comprises a process for manufacturing an intermediatelayer between an internal jacket and an external jacket of a combustionchamber, in particular a rocket drive, wherein at least one part of thestructure is produced from a composite material on the basis of afibrous structure made of carbon-containing fibers. First, a fibrousstructure made of carbon-containing fibers is affixed on the internaljacket, and then a metal material is deposited on the fibrous structurewith simultaneous infiltration of the fibrous structure with the metalmaterial. The metal material can be deposited in any suitable form, forexample, in a liquid or gas phase. However, the metal material can alsobe deposited preferably by way of an electroplating process. Inparticular, nickel or copper or a nickel- and/or copper-containing alloycan be used here as the metal material.

Other objects, advantages and novel features of the present inventionwill become apparent from the following detailed description of theinvention when considered in conjunction with the accompanying drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 a shows a fibrous structure with layers of fibers which extend intwo directions in space;

FIG. 1 b shows the fibrous structure of FIG. 1 a, interwoven withanadditional layer of fibers;

FIG. 2 a shows a detailed view of the wall structure of a rocketcombustion chamber with an intermediate layer and regenerative cooling;

FIG. 2 b shows a sectional view of a rocket combustion chamber accordingto FIG. 2 a; and

FIG. 3 shows a detailed view of the wall structure of an alternativerocket combustion chamber without intermediate layer but withregenerative cooling.

DETAILED DESCRIPTION OF THE DRAWINGS

With the aid of textile methods, such as interweaving or sewing, afibrous structure is produced in the form of a three-dimensional matrixmade of carbon fibers and then infiltrated with silicon, forming asilicon carbide-containing composite material. FIGS. 1 a and 1 b depicta corresponding fibrous structure. They show several layers 1 a, 1 b, 1c of fibers or bundles of fibers 10, wherein the fibers or bundles offibers of the layer 1 a extend in the y direction, those of layer 1 bextend in the x direction, and those of the layer 1 c extend in the zdirection. Thus, the individual layers and/or the fibers or the bundlesof fibers 10 of the individual layers extend in different directions inspace. In addition, several fibers or bundles of fibers 10 are combinedto form larger bundles 11, wherein these larger bundles 11 of a layer 1a, 1 b, 1 c are spaced apart. The amount of spacing is chosen in such amanner that at least one larger bundle 11 of another layer 1 a, 1 b, 1 ccan penetrate at this point the respective layer 1 a, 1 b, 1 c. In thismanner, the individual layers 1 a, 1 b, 1 c are interwoven. Thissituation is depicted in FIG. 1 b. As an alternative or in addition, theindividual layers 1 a, 1 b, 1 c, or the individual fibers or bundles offibers 10, or the larger bundles 11, can be sewn together.

The provision of layers 1 a, 1 b, 1 c, whose fibers or bundles of fibers10 extend in three directions in space x, y, z, results in a defined,three-dimensional fiber matrix, which guarantees that the subsequentcomposite material is stable against shear forces, generated, forexample, by thrust loads in rocket drives. A two-dimensional fibrousstructure does not exhibit this strength. Rather, high shear forces canresult in interlaminar failure in a two-dimensional fibrous structure.Thus, one significant advantage of the defined three-dimensional fibrousstructure lies in increased tolerance. In addition, with theintroduction of special types of fibers and special matrix systems, themechanical and physical properties of the subsequent composite material,such as thermal conductivity and expansion coefficient, can be adjustedas desired.

An example of combustion chamber structure cooled by transpiration isshown in FIG. 2 a. In such a structure and at a selected porosity, forexample, at a porosity ranging from 10% to 30%, such as approximately20%, the three-dimensional fibrous structure guarantees adequate bondingstrength or damage tolerance, as compared to the two-dimensionalconstruction, particularly for components with stresses in severaldirections in space.

Cooling channels 5 a can be placed in the combustion chamber structure.Cooling channels 5 a may be especially useful in rocket driveapplications. Cooling channels 5 a can be placed either with the usedcontracting bodies, which dissolve or decompose in the process, forexample, plastics or wax, also referred to as lost cores because theycan be dissolved out thermally by melting or pyrolysis. The contractingbodies are placed in the fibrous structure prior to fabrication of thecomposite material or they are placed in the channels of the compositematerial. Then the channels are closed with a cover layer. However, thechannels 5 a can also be worked mechanically into the finished compositematerial, for example, by milling. Channels 5 a can also be manufacturedby in-situ manufacture via textile technology, for example, byinterweaving with simultaneous release of the correspondingchannel-shaped openings.

The result of placing special intermediate materials in the form of anintermediate layer 7 a between the ceramic combustion chamber compositelayer as the internal jacket 4 a and a metal, load-bearing externalstructure as the external jacket 8 a is that any stresses in theconnecting zone (metal/ceramic) can be reduced to a subcritical amount.In this respect the intermediate layer 7 a is designed in such a mannerthat it exhibits a defined expansion coefficient between that of thematerial of the internal jacket 4 a and that of the external jacket 8 a(graduated construction). For example, a metal matrix composite (MMC), acomposite material with a metal matrix, can be used as the intermediatematerial. This intermediate material can be produced, for example, byaffixing a fibrous structure made of carbon-containing fibers on theinternal jacket. The fibrous structure may be made of silicon carbide orcarbon, in the form of a matrix, which, like that of the internaljacket, can be interwoven or braided, or as a fleece or wound fibers,which can be long or short. This fibrous structure can then beinfiltrated with a metal material, like nickel, copper, or an alloy withnickel and/or copper in accordance with an electroplating process,wherein at the same time an electroplated layer for forming the externaljacket 8 a is produced on the fibrous structure. Exclusivelycarbon-containing particles can also be used to form the MMC. The MMCcan also be produced by way of the methods known in the art. Then themetal external structure 8 a can be affixed, for example, by soldering,such as active soldering, or electroplating technology.

Many variants for manufacturing combustion chambers according to thepresent invention exist, certain methods are described in detail below.

EXAMPLE 1 A Drive with Regenerative Cooling

This example may be best understood by referring to FIGS. 2 a and 2 b.First, the manufacture of the two-dimensional layer construction 1 a, 1b is accomplished by winding, prepreg, or weaving technology from fibersor bundles of fibers 10. The third fiber layer 1 c is incorporated bytextile technology, such as tufting, sewing, and weaving. The structuremay optionally be hardened for shape stability in the carbon fiberreinforced plastic state. Segmentation (slots) 3 a of the combustionchamber structure are optionally provided for minimizing stress.

The multiaxial structure is then pyrolyzed/carbonized to form a denseC/SiC or C/C structure 4 a by repeated post-infiltration or Siinfiltration. Cooling channels 5 a are placed by mechanical treatment ofthe composite material, by textile technology, or by contracting bodiesprior to completion of the composite material. The cooling channels areoptionally sealed with a metal coating 6 a for improved pressureresistance and tightness with respect to the coolant that flows throughin the operating state. Cooling channels may optionally be reworded bymilling.

Contracting bodies are placed in the structure, as appropriate. For theelectroplating variant, the contracting bodies may be placed byelectrically conducting wax into the channels 5 a. A metal intermediatelayer 7 a is then affixed. This may be accomplished, for example, byaffixing MMC with a defined expansion coefficient (AK) between thematerial of the internal jacket 4 a (ceramic composite) and the externaljacket 8 a (metal). After affixing the metal intermediate layer, themetal external structure (8 a) is affixed, for example, byelectroplating or by active soldering technology such as using thematerial Inconel 718, a nickel-based alloy. Finally, the contractingbodies are dissolved, for example, by melting out the wax.

EXAMPLE 2 A Drive with Transpirative Cooling, thus with a PorousCombustion Chamber Structure

This example may be best understood by reference to FIGS. 2 a and 2 b.The two-dimensional layer construction 1 a, 1 b is manufactured bywinding, prepreg, or weaving technology with simultaneous placement ofplace-holders (doped precursors) as the contracting bodies. The thirdfiber layer 1 c is incorporated by textile technology such as tufting,sewing, or weaving. The structure may optionally be hardened for shapestability in the carbon fiber reinforced plastic state. Segmentation(slots) 3 a of the combustion chamber structure can be optionallyprovided for minimizing stress.

The multiaxial structure is pyrolyzed/carbonized to form a porous (openpores) C/SiC or C/C structure 4 a with pores or micro cracks. At thesame time, the place-holders are burned out by pyrolysis, during whichthe porosity of the structure can be adjusted by manipulating the degreeof infiltration, for example, of silicon. Cooling channels 5 a mayoptionally be added by mechanical treatment or by contracting bodies.The contracting bodies can be placed, for example, by electricallyconducting wax into the channels for electroplating variants.

A metal intermediate layer 7 a may optionally be affixed, for example,MMC with a defined expansion coefficient (AK) may be affixed between thematerial of the internal jacket 4 a (ceramic composite) and the externaljacket 8 a (metal). The metal external structure 8 a is affixed byelectroplating, such as with nickel, or by active soldering technology,for example, using the material Inconel 718, a nickel-based alloy.Finally, the contracting bodies are dissolved, for example, by meltingout the wax.

EXAMPLE 3 A Drive with Regenerative Cooling

This example may be best understood by reference to FIG. 3. IndividualU-profiles 4 c made of a fiber composite ceramic with a definedthree-dimensional fibrous structure or Si/SiC analogous aremanufactured, such as according to Examples 1 or 2. The U-profiles 4 care bundled/fixed. Optionally, the surface coating 2 c may beelectroplated, for example, with copper or nickel, at least on theinside of the U-profiles 4 c. This provides improved pressure resistanceand tightness with respect to the coolant flowing through in theoperating state. Then the U-profile coating 2 c is optionally reworkedand the channel-shaped U-profile structures 5 a are filled withcontracting bodies, for example, conductive wax. The metal externalstructure 3 c is deposited by electroplating with material such asnickel. Finally, the wax is melted out.

The foregoing disclosure has been set forth merely to illustrate theinvention and is not intended to be limiting. Since modifications of thedisclosed embodiments incorporating the spirit and substance of theinvention may occur to persons skilled in the art, the invention shouldbe construed to include everything within the scope of the appendedclaims and equivalents thereof.

1. A combustion chamber for a rocket drive, comprising: at least onefirst jacket made of a composite material with a ceramic matrix; and aload bearing external jacket that envelopes said at least one firstjacket; wherein, said composite material comprises a fibrous structuremade of carbon-containing fibers; and said fibrous structure comprisesfirst, second and third layers of fiber bundles forming athree-dimensional matrix; fiber bundles of said first layer of fiberbundles extend in a first direction in space; fiber bundles of saidsecond layer of fiber bundles extend in a second direction in space;fiber bundles of the third layer of fiber bundles extend in a thirddirection in space, said first, second and third directions beingsubstantially divergent relative to each other; and the individuallayers penetrate each other at least partially.
 2. A combustion chamberaccording to claim 1, wherein said second layer and said third layer ofsaid fibrous structure are connected by means of textile technology. 3.A combustion chamber according to claim 1, wherein said fibrousstructure comprises carbon fibers.
 4. A combustion chamber according toclaim 1, wherein said composite material comprises silicon carbide.
 5. Acombustion chamber according to claim 1, wherein said external jacket ismade of metal.
 6. A combustion chamber according to claim 5, whereinsaid metal matrix contains the same metal material as said externaljacket.
 7. A combustion chamber for a rocket drive, comprising: at leastone first jacket made of a composite material with a ceramic matrix; aload bearing external jacket that envelopes said at least one firstjacket; and an intermediate layer disposed between said external jacketand said at least one first jacket; wherein, said composite material ofsaid at least one first jacket comprises a fibrous structure made ofcarbon-containing fibers; and said fibrous structure comprises layers offibers forming a three-dimensional material; said intermediate layercomprises a composite material with a metal matrix; and a thermalexpansion coefficient of said intermediate layer is between that of saidexternal jacket and said at least one first jacket.